EP1508671A2 - Joint à brosses pour une turbomachine - Google Patents

Joint à brosses pour une turbomachine Download PDF

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Publication number
EP1508671A2
EP1508671A2 EP04026038A EP04026038A EP1508671A2 EP 1508671 A2 EP1508671 A2 EP 1508671A2 EP 04026038 A EP04026038 A EP 04026038A EP 04026038 A EP04026038 A EP 04026038A EP 1508671 A2 EP1508671 A2 EP 1508671A2
Authority
EP
European Patent Office
Prior art keywords
gas turbine
blade outer
seal
outer air
turbine engine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP04026038A
Other languages
German (de)
English (en)
Other versions
EP1508671B1 (fr
EP1508671A3 (fr
Inventor
Joseph P. Bouchard
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP1508671A2 publication Critical patent/EP1508671A2/fr
Publication of EP1508671A3 publication Critical patent/EP1508671A3/fr
Application granted granted Critical
Publication of EP1508671B1 publication Critical patent/EP1508671B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16J—PISTONS; CYLINDERS; SEALINGS
    • F16J15/00—Sealings
    • F16J15/16—Sealings between relatively-moving surfaces
    • F16J15/32—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings
    • F16J15/3284—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings characterised by their structure; Selection of materials
    • F16J15/3288—Filamentary structures, e.g. brush seals
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/55—Seals
    • F05D2240/56—Brush seals

Definitions

  • the present invention relates to gas turbine engines and, more particularly, to a brush seal for minimizing air leakage therein.
  • a typical gas turbine engine includes a low compressor, a high compressor, a combustor, a high turbine, and a low turbine sequentially situated about a longitudinal axis.
  • the sections of the gas turbine engine are enclosed in an engine case with air flowing axially through the engine along an annular air path.
  • the air enters the low compressor at a substantially ambient temperature and pressure and is compressed by the low and high compressor sections, respectively, to elevate its temperature and pressure.
  • the compressed air is then mixed with fuel, ignited, and burned in the combustor.
  • the hot products of combustion emerging from the combustor are expanded in the high and low turbines, respectively, thereby rotating the turbine and driving the compressor.
  • the high turbine section of the gas turbine engine is subjected to extremely harsh environment, characterized by very high temperatures and pressures.
  • the components of the high pressure turbine must be cooled to prevent these components from burning in a very short period of time.
  • the cooling air is typically bled from the high pressure compressor and routed to the high pressure turbine. Although the bled cooling air is necessary to cool certain engine parts, loss of the cooling air from the compressor is highly undesirable.
  • the cooling air, diverted from the compressor is no longer available to produce thrust and thus negatively effects the efficiency of the gas turbine engine. Therefore, the cooling air is carefully routed from the high pressure compressor to the high pressure turbine to ensure that this valuable cooling air is not wasted.
  • the cooling air is routed through the engine into cooling air compartments that are substantially adjacent to the high pressure turbine air path.
  • the cooling air from the cooling air compartments is carefully metered out to cool various components of the turbine section of the gas turbine engine.
  • a barrier is required to prevent the cooling air from escaping from the cooling air compartment into the turbine air path, other than through designated bleeding holes, since the cooling air is at a greater pressure than the air in the turbine air path.
  • the conventional barriers are annular metal seals which are not completely effective for a number of reasons.
  • certain sealing surface which abut these seals are not uniform.
  • a plurality of blade outer air seals forms a segmented ring around the tips of rotating blades.
  • Even slight size differences among the blade outer air seals results in gaps between some of the blade outer air seals and the barrier seals, thereby allowing some of the cooling air to escape therethrough.
  • a second difficulty with effectively sealing the cooling air compartments located in the high pressure turbine of the gas turbine engine is that the high pressure turbine experiences great temperature fluctuations. As a result of these large temperature gradients, the turbine components are subjected to significant thermal expansion. When components of the high pressure turbine expand, the conventional seals cannot adjust and seal the gap between expanded components and the seal, therefore, allowing some of the cooling air to escape through these gaps. Therefore, there is a great need to effectively seal cooling air compartments in the high pressure turbine.
  • a brush seal for a gas turbine engine includes a first ring and a second ring with a plurality of bristles sandwiched therebetween and extending in an axial direction to minimize high pressure cooling air leakage into an air path of the gas turbine engine.
  • the brush seal extends axially between a vane support and a plurality of blade outer air seals with the bristle tips compressed against a sealing surface of the vane support and against sealing surfaces of the blade outer air seals.
  • the brush seal rests against a plurality of flanges formed on each of the blade outer air seals such that the movement of the brush seal is not significantly restricted.
  • the axial brush seal maintains contact with both sealing surfaces. Additionally, the axial brush seal compensates for any lack of uniformity in the blade outer air seals by having some of the bristle tips slightly more compressed or slightly less compressed.
  • an axial brush seal embodying the present invention can be used to seal components that do not have a uniform circumferential surface and to compensate for the thermal growth of components in high temperature sections of the gas turbine engine.
  • the brush seal includes a seal lip protruding outward from the first ring of the brush seal for properly orienting the brush seal within the gas turbine engine.
  • the seal lip also provides an additional barrier against air leakage.
  • One advantage of the present invention in its preferred embodiments is that the seal does not need to be bolted or otherwise fixedly secured, thereby reducing the cost of the seal and reducing assembly time during manufacturing. Also, placing bolts or rivets through the seal degrades the quality of the seal and results in loose bristles that tend to fall out from the seal, thereby reducing effectiveness of the seal.
  • a gas turbine engine 10 includes a compressor 12, a combustor 14, and a turbine 16 sequentially situated about a longitudinal axis 18.
  • a gas turbine engine case 20 encloses sections 12, 14, 16 of the gas turbine engine 10. Air flows substantially axially along an annular air path 22 through the sections 12, 14, 16 of the gas turbine engine 10 and is exhausted through an aft portion 24 of the gas turbine engine 10.
  • the compressor 12 comprises a low compressor section 26 and a high compressor section 28 and the turbine 16 comprises a high turbine section 32 and a low turbine section 34.
  • the compressor 12 and the turbine 16 include alternating rows of stationary vanes 36 and rotating blades 38.
  • the stationary vanes 36 are mounted onto the engine case 20.
  • the rotating blades 38 of the low compressor 26 and low turbine 34 are secured onto a rotating low rotor 42.
  • the rotating blades 38 of the high compressor 28 and high turbine 32 are secured onto a rotating high rotor 44, which is disposed radially outward of the rotating low rotor 42.
  • the high pressure turbine section 32 of the gas turbine engine 10 includes substantially annular first and second vane supports 48, 50 secured to the engine case 20 for supporting vane platform portions 52 of the vanes 36.
  • the first vane support 48 includes a vane support sealing surface 53.
  • the high pressure turbine 32 also includes a substantially annular blade outer air seal support 54 secured to the engine case 20 and sandwiched between the first and second vane supports 48, 50.
  • a first air cavity 56 and a second air cavity 58 are defined between the blade outer air seal support 54 and the vane support 48.
  • the blade outer air seal support 54 supports a plurality of blade outer air seals 62 positioned radially outward of the rotating blades 38.
  • the plurality of blade outer air seals 62 forms a segmented ring around the outer circumference of the air path 22 to define a running clearance 64 between the blade outer air seals 62 and blade tips 66 of the rotating blades 38.
  • Each blade outer air seal 62 includes blade outer air seal first and second tabs 70, 72 that fit into the blade outer air seal support 54, defining a third cooling air cavity 74 therebetween.
  • Each blade outer air seal 62 also includes a blade outer air seal sealing surface 76 and a blade outer air seal clearance surface 78.
  • Each of the blade outer air seals 62 further includes a plurality of blade outer air seal cooling openings 82 allowing cooling air from the third cooling air cavity 74 to pass through to cool the clearance surface 78 and a flange 83 protruding radially outward.
  • the high pressure turbine 32 further includes an axial brush seal 84 extending axially between the sealing surface 53 of the vane support 48 and the sealing surfaces 76 of the blade outer air seals 62.
  • the axial seal 84 is supported by the flanges 83 of the blade outer air seals 62.
  • the axial brush seal 84 includes a first seal ring 86 and a second seal ring 88 with a plurality of bristles 92 extending substantially axially therebetween.
  • the first ring 86 includes an axial seal lip 94 protruding radially outward to fit into a blade outer air support groove 96, as best seen in FIG. 2.
  • Each of the plurality of bristles 92 has a fore bristle end 98 and an aft bristle end 100.
  • the first and second rings 86, 88 with a plurality of bristles 92 are welded together in the center thereof to form a substantially circumferential seam 102.
  • the pluralities of bristles 92 extend from the first and second rings 86, 88 at an angle, as shown in FIG. 4.
  • the pressure of the air flowing along the airpath 22 sequentially through the low and high compressor 26, 28 is increased, thereby effecting compression of the incoming airflow 22, as best seen in FIG. 1.
  • a portion of the compressed air is bled and routed to the high pressure turbine 32.
  • the remaining compressed air is mixed with fuel, ignited and burned in the combustor 14.
  • the hot products of combustion emerge from the combustor 14 and enter the high turbine 32 at extremely high temperature and pressure.
  • the turbine blades 38 expand the hot air, generating thrust and extracting energy to drive the compressor 12.
  • the axial brush seal 84 minimizes leakage of the cooling air from the cooling air cavities 56, 58, 74 into the air path 22.
  • the bristles 92 of the axial seal 84 are compressed between the vane support 48 and the blade outer air seal 62 with the plurality of aft bristle tips 100 of the axial brush seal 84 being compressed against the blade outer air seal sealing surface 76 and the plurality of fore bristle tips 98 of the brush seal 84 being compressed against the vane support sealing surface 53.
  • Some of the aft bristle tips 100 are slightly more compressed and some of the aft bristle tips 100 are slightly less compressed to compensate for lack of uniformity in size and manufacturing variations in the blade outer air seals 62.
  • the rotating blades 38 experience thermal expansion. Also, as a result of centrifugal forces, the tips 66 of the rotating blades 38 move radially outward toward the blade outer air seals 62. In order to maintain the blade tip clearance 64 small throughout all operational phases of gas turbine engine 10, but to avoid interference between the blade tips 66 and blade outer air seals 62, the blade outer air seals 62 need to move radially outward as the blade tips 66 expand radially outward.
  • the growth of the blade outer air seal support 54 and of the blade outer air seals 62 is precisely controlled by routing the cooling air into the cooling air cavities 56, 58, 74 and by metering out cooling air through the cooling openings 82 of the blade outer air seals 62.
  • the vane support 48 moves radially and axially as the result of thermal expansion.
  • the plurality of fore bristle tips 98 continues to be in contact with the vane support sealing surface 53, either further compressing or further expanding.
  • the axial seal 84 moves radially with the thermally growing or shrinking blade outer air seals 62 as the axial seal 84 rests on the flanges 83 of the blade outer air seals 62.
  • the seal lip 94 allows some radial movement of the axial seal 84 and functions as a further air flow deterrent for axial air leakage.
  • the seal lip 94 ensures proper placement of the axial seal 84 in the gas turbine engine 10.
  • the axial brush seal 84 positioned between the vane support 48 and plurality of blade outer air seals 62, compensates for thermal expansion of adjacent components as well as for size and dimensional variations of the blade outer air seal segments 62.
  • One advantage of the present invention is that the axial seal 84 does not need to be bolted or otherwise fixedly secured, thereby reducing the cost of the seal and reducing assembly time during manufacturing. Also, placing bolts or rivets through the seal degrades the quality of the seal and results in loose bristles that tend to fall out from the seal, thereby reducing the seal's effectiveness.
  • the bristles 92 extend from the seal rings 86, 88 at an angle, as shown in FIG. 4.
  • the preferred angle is approximately forty-five degrees (45°).
  • bristles 92 can extend substantially perpendicular to the seal rings 86, 88 or form other angles therewith.
  • the seal lip 94 is not essential for certain applications where orientation of the axial seal 84 is not critical.
  • a floating brush seal that is not fixedly attached to any components can also extend in the radial direction when air leaks in the axial direction.

Landscapes

  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Sealing Devices (AREA)
EP04026038A 1998-12-23 1999-12-15 Joint à brosses pour une turbomachine Expired - Lifetime EP1508671B1 (fr)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US221289 1998-12-23
US09/221,289 US6170831B1 (en) 1998-12-23 1998-12-23 Axial brush seal for gas turbine engines
EP99310104A EP1013887B1 (fr) 1998-12-23 1999-12-15 Joint à brosses pour une turbomachine

Related Parent Applications (1)

Application Number Title Priority Date Filing Date
EP99310104A Division EP1013887B1 (fr) 1998-12-23 1999-12-15 Joint à brosses pour une turbomachine

Publications (3)

Publication Number Publication Date
EP1508671A2 true EP1508671A2 (fr) 2005-02-23
EP1508671A3 EP1508671A3 (fr) 2005-03-23
EP1508671B1 EP1508671B1 (fr) 2007-11-28

Family

ID=22827176

Family Applications (2)

Application Number Title Priority Date Filing Date
EP99310104A Expired - Lifetime EP1013887B1 (fr) 1998-12-23 1999-12-15 Joint à brosses pour une turbomachine
EP04026038A Expired - Lifetime EP1508671B1 (fr) 1998-12-23 1999-12-15 Joint à brosses pour une turbomachine

Family Applications Before (1)

Application Number Title Priority Date Filing Date
EP99310104A Expired - Lifetime EP1013887B1 (fr) 1998-12-23 1999-12-15 Joint à brosses pour une turbomachine

Country Status (4)

Country Link
US (1) US6170831B1 (fr)
EP (2) EP1013887B1 (fr)
JP (1) JP2000193095A (fr)
DE (2) DE69926332T2 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8794918B2 (en) 2011-01-07 2014-08-05 General Electric Company System for adjusting brush seal segments in turbomachine
US9121297B2 (en) 2011-03-28 2015-09-01 General Electric Company Rotating brush seal
US9255486B2 (en) 2011-03-28 2016-02-09 General Electric Company Rotating brush seal
EP2670954B1 (fr) * 2011-01-31 2020-05-06 General Electric Company Joint à brosse axial

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Publication number Priority date Publication date Assignee Title
US20100007093A1 (en) * 2001-02-23 2010-01-14 Grondahl Clayton M Seal Assembly and Rotary Machine Containing Such Seal
US7578509B2 (en) * 2001-02-23 2009-08-25 Cmg Tech, Llc Seal assembly and rotary machine containing such seal
US6644667B2 (en) 2001-02-23 2003-11-11 Cmg Tech, Llc Seal assembly and rotary machine containing such seal
US6622490B2 (en) * 2002-01-11 2003-09-23 Watson Cogeneration Company Turbine power plant having an axially loaded floating brush seal
US7093835B2 (en) * 2002-08-27 2006-08-22 United Technologies Corporation Floating brush seal assembly
US6984106B2 (en) * 2004-01-08 2006-01-10 General Electric Company Resilent seal on leading edge of turbine inner shroud
FR2865012B1 (fr) * 2004-01-12 2006-03-17 Snecma Moteurs Dispositif d'etancheite pour turbine haute-pression de turbomachine
US7721433B2 (en) * 2005-03-28 2010-05-25 United Technologies Corporation Blade outer seal assembly
US20060267291A1 (en) * 2005-05-25 2006-11-30 Mark Addis Brush seal assembly and method of assembling same
US20070132190A1 (en) * 2005-12-12 2007-06-14 Charles Trabert Axial dynamic brush seal
US7665955B2 (en) * 2006-08-17 2010-02-23 Siemens Energy, Inc. Vortex cooled turbine blade outer air seal for a turbine engine
US8414254B2 (en) * 2008-02-18 2013-04-09 United Technologies Corporation Sealing assembly for a turbine engine
US8596973B2 (en) * 2009-12-07 2013-12-03 Cmg Tech, Llc Leaf seal assembly including polymer member and rotary machine containing such seal assembly
US8632075B2 (en) * 2011-08-08 2014-01-21 General Electric Company Seal assembly and method for flowing hot gas in a turbine
US9316119B2 (en) 2011-09-15 2016-04-19 United Technologies Corporation Turbomachine secondary seal assembly
US20130170979A1 (en) * 2012-01-04 2013-07-04 General Electric Company Double ended brush seal assembly for a compressor
WO2013153866A1 (fr) * 2012-04-08 2013-10-17 イーグル工業株式会社 Joint brosse
US9574455B2 (en) * 2012-07-16 2017-02-21 United Technologies Corporation Blade outer air seal with cooling features
FR2998611B1 (fr) 2012-11-29 2018-08-10 Safran Aircraft Engines Ensemble formant joint d'etancheite pour une turbomachine comportant un joint a brosse
US10208674B2 (en) 2013-03-13 2019-02-19 United Technologies Corporation Multi-axial brush seal
WO2014146866A1 (fr) * 2013-03-20 2014-09-25 Siemens Aktiengesellschaft Élément d'étanchéité permettant d'assurer l'étanchéité d'une fente et turbine à gaz correspondante
FR3005487B1 (fr) 2013-05-13 2015-06-05 Snecma Ensemble formant joint d'etancheite pour une turbomachine comportant des moyens de lubrification d'un joint a brosse
EP2843196B1 (fr) * 2013-09-03 2020-04-15 Safran Aero Boosters SA Compresseur de turbomachine et turboachine associée
EP3044427B8 (fr) 2013-09-12 2021-04-07 Raytheon Technologies Corporation Moteur à turbine à gaz et procédé de réglage de jeu d'extrémités d'aubes
WO2015080779A2 (fr) * 2013-09-13 2015-06-04 United Technologies Corporation Joint d'étanchéité haute température à déplacement important
WO2015116399A1 (fr) * 2014-01-28 2015-08-06 United Technologies Corporation Joint de cavité flexible pour moteurs à turbine à gaz
US9879557B2 (en) * 2014-08-15 2018-01-30 United Technologies Corporation Inner stage turbine seal for gas turbine engine
US10400896B2 (en) 2014-08-28 2019-09-03 United Technologies Corporation Dual-ended brush seal assembly and method of manufacture
US9896955B2 (en) * 2015-04-13 2018-02-20 United Technologies Corporation Static axial brush seal with dual bristle packs
US9863538B2 (en) 2015-04-27 2018-01-09 United Technologies Corporation Gas turbine engine brush seal with supported tip
US9869195B2 (en) 2015-05-19 2018-01-16 United Technologies Corporation Support assembly for a gas turbine engine
US9885247B2 (en) 2015-05-19 2018-02-06 United Technologies Corporation Support assembly for a gas turbine engine
US9896956B2 (en) * 2015-05-22 2018-02-20 United Technologies Corporation Support assembly for a gas turbine engine
US10385712B2 (en) 2015-05-22 2019-08-20 United Technologies Corporation Support assembly for a gas turbine engine
US10107128B2 (en) 2015-08-20 2018-10-23 United Technologies Corporation Cooling channels for gas turbine engine component
US10428953B2 (en) * 2016-02-25 2019-10-01 United Technologies Corporation C-seal backed brush seal with a compressible core
US10208614B2 (en) * 2016-02-26 2019-02-19 General Electric Company Apparatus, turbine nozzle and turbine shroud
US9845884B2 (en) 2016-03-17 2017-12-19 United Technologies Corporation Brush seal with single-layer mixed-diameter bristle pack
US20170292395A1 (en) * 2016-04-07 2017-10-12 United Technologies Corporation Integrated brush seals
US10563771B2 (en) * 2016-04-07 2020-02-18 United Technologies Corporation Wire mesh brush seal windage cover
US10352183B2 (en) * 2016-04-25 2019-07-16 United Technologies Corporation High temperature seal and method
US10280799B2 (en) * 2016-06-10 2019-05-07 United Technologies Corporation Blade outer air seal assembly with positioning feature for gas turbine engine
BE1025092B1 (fr) * 2017-03-31 2018-10-29 Safran Aero Boosters S.A. Joint a brosse pour rotor de turbomachine
US10502093B2 (en) * 2017-12-13 2019-12-10 Pratt & Whitney Canada Corp. Turbine shroud cooling
US10633995B2 (en) 2018-07-31 2020-04-28 United Technologies Corporation Sealing surface for ceramic matrix composite blade outer air seal
US10787923B2 (en) * 2018-08-27 2020-09-29 Raytheon Technologies Corporation Axially preloaded seal
US10968761B2 (en) * 2018-11-08 2021-04-06 Raytheon Technologies Corporation Seal assembly with impingement seal plate
US10753221B2 (en) 2018-12-12 2020-08-25 Raytheon Technologies Corporation Seal assembly with ductile wear liner
US11015473B2 (en) * 2019-03-18 2021-05-25 Raytheon Technologies Corporation Carrier for blade outer air seal
EP4438764A1 (fr) 2023-03-29 2024-10-02 Pratt & Whitney Canada Corp. Revêtement à plusieurs composants

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EP0534686A1 (fr) * 1991-09-23 1993-03-31 General Electric Company Joint-brosse disposé dans une gorge
GB2301635A (en) * 1995-04-12 1996-12-11 Gec Alsthom Ltd Shaft seal arrangement
EP0775805A2 (fr) * 1995-11-22 1997-05-28 United Technologies Corporation Anneau de stator

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8794918B2 (en) 2011-01-07 2014-08-05 General Electric Company System for adjusting brush seal segments in turbomachine
EP2670954B1 (fr) * 2011-01-31 2020-05-06 General Electric Company Joint à brosse axial
US9121297B2 (en) 2011-03-28 2015-09-01 General Electric Company Rotating brush seal
US9255486B2 (en) 2011-03-28 2016-02-09 General Electric Company Rotating brush seal

Also Published As

Publication number Publication date
US6170831B1 (en) 2001-01-09
DE69937652D1 (de) 2008-01-10
DE69926332T2 (de) 2006-05-24
EP1508671B1 (fr) 2007-11-28
EP1508671A3 (fr) 2005-03-23
EP1013887B1 (fr) 2005-07-27
DE69926332D1 (de) 2005-09-01
DE69937652T2 (de) 2008-10-30
EP1013887A1 (fr) 2000-06-28
JP2000193095A (ja) 2000-07-14

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